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WO2014031275A1 - Éléments de refroidissement interne de surface portante de moteur à turbine à gaz - Google Patents

Éléments de refroidissement interne de surface portante de moteur à turbine à gaz Download PDF

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Publication number
WO2014031275A1
WO2014031275A1 PCT/US2013/051796 US2013051796W WO2014031275A1 WO 2014031275 A1 WO2014031275 A1 WO 2014031275A1 US 2013051796 W US2013051796 W US 2013051796W WO 2014031275 A1 WO2014031275 A1 WO 2014031275A1
Authority
WO
WIPO (PCT)
Prior art keywords
span
airfoil according
airfoil
trip
pressure
Prior art date
Application number
PCT/US2013/051796
Other languages
English (en)
Inventor
Daniel C. NADEAU
Jeffrey R. Levine
Dominic J. Mongillo Jr.
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to EP13830772.3A priority Critical patent/EP2888462B1/fr
Publication of WO2014031275A1 publication Critical patent/WO2014031275A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/127—Vortex generators, turbulators, or the like, for mixing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/305—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the pressure side of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/306—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the suction side of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This disclosure relates to a gas turbine engine airfoil.
  • the disclosure relates to the airfoil's internal cooling features.
  • cooling passageways extend radially within the airfoil.
  • Three discrete feed cavities are provided in the root.
  • a leading edge feed cavity includes two cooling passageways, a midbody feed cavity includes three cooling passageways, and a trailing edge cavity includes three cooling passageways. Only skewed trip strips are provided on the walls defining the cooling passageways to transfer heat to the cooling fluid flowing through the cooling passageways.
  • the trip strip layout uses skewed trip strips from 10%-30% span, and no trip strips from 0%-10% span. From 10% to 30% span, the trip strip height is 0.010 inch (0.25mm) with a P/E (trip radial pitch to trip height) of 10. Pitch is the spacing between adjacent trip strips. From 30% to 100% span, the trip strip height of the skewed trip strips is 0.015 inch (0.38mm) and the P/E is 5 for the leading edge and midbody cooling passageways. For the trailing edge cooling passageways, the trip strip height from 30%-!0Q% span is 0.012 inch (0.30mm) with a P/E of 6.25.
  • an airfoil for a gas turbine engine includes spaced apart pressure and suction walls joined at leading and trailing edges to provide an airfoil. Intermediate walls interconnect the pressure and suction walls to provide cooling passageways.
  • the cooing passageways have interior pressure and suction surfaces that are respectively provided on the pressure and suction walls.
  • Trip strips include a chevron-shaped trip strip that is provided on at least one of the interior pressure and suction surfaces.
  • the airfoil extends in a radial direction from a platform supported on a root to a tip.
  • the cooling passageways extend in the radial direction.
  • the chevron-shaped trip strip includes protrusions joined at an apex. The apex faces a flow direction within the cooling passageway.
  • feed cavities are provided in the root and are fluidly connected to the cooling passageways.
  • Each feed cavity has an inlet discrete from the other feed cavities.
  • the cooling passageways are spaced apart from one another in a chord- wise direction that extends between the leading and trailing edges.
  • the feed cavities correspond to leading edge, midbody and trailing edge feed cavities.
  • the leading edge feed cavity has first and second cooling passageways.
  • the midbody feed cavity has third, fourth and fifth cooling passageways.
  • the trailing edge feed cavity has sixth, seventh and eighth cooling passageways.
  • the airfoil includes first, second and third span regions extending along a span that extends in the radial direction from 0% span near the platform to 100% span near the tip.
  • the first span region extends 0-20% the span +1-5%.
  • the second span region extends 20-60% the span +1-5%.
  • the third span region extends 60- 100% the span +1-5%.
  • the first, second and third span regions have different average P/E ratios, wherein P is a pitch of the trip strips, and E is a height of the trip strip with respect to a support surface of the trip strip.
  • the height of trip strips within the first and second span regions in the leading edge and midbody feed cavities is 0.015 inch (0.38 mm) +/-0.002 inch (0.05 mm).
  • the first span region in the leading edge and midbody feed cavities has an average P/E ratio of 6.7 +/-0.3.
  • the second span region in the leading edge and midbody feed cavities has an average P/E ratio of 5.0 +/-0.3.
  • the height of trip strips within the first and second span regions in the trailing edge feed cavity is 0.012 inch (0.30 mm) +/-0.002 inch (0.05 mm).
  • the first span region in the trailing edge feed cavity has an average P/E ratio of 8.3 +/-0.3.
  • the second span region in the trailing edge feed cavity has an average P/E ratio of 6.3 +/-0.3.
  • the trip strips include single linear trip strips skewed relative to the flow direction.
  • the chevron trip strips extend from 0% to 60% span +/-5% in the second and sixth cooling passageways.
  • the chevron trip strips extend from 0% to 100% span +1-5% in the fifth cooling passageway.
  • the from 0% to 20% span +/- 5% includes no trip strips on the suction surface.
  • the from 0% to 20% span +/- 5% includes trip strips on the interior pressure surface.
  • the pressure and suction surfaces include different trip strip configurations.
  • the different trip strip configuration includes skewed and chevron features.
  • the protrusions are of unequal length.
  • Figure 1 is a schematic cross-sectional view of an example gas turbine engine.
  • Figure 2 is a perspective view of an example turbine blade.
  • Figure 3 is a schematic cross-sectional view through the blade shown in Figure 2 taken along line 3-3.
  • Figure 4 is a cross-sectional view taken through the airfoil in Figure 2 along line 4-4.
  • Figure 5 is a view of the airfoil cooling passageways illustrating the internal cooling features on a pressure side of the airfoil.
  • Figure 6 is a view of the airfoil cooling passageways illustrating the internal cooling features on a suction side of the airfoil.
  • Figure 7 is a perspective view of an example chevron-shaped trip strip on an interior surface of a cooling passageway of the airfoil.
  • FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmenter section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26.
  • air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
  • turbofan gas turbine engine depicts a turbofan gas turbine engine
  • the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
  • the example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46.
  • the inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54.
  • the high pressure turbine 54 includes only a single stage.
  • a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure” compressor or turbine.
  • the core airflow C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5: 1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet (10,668 m).
  • the flight condition of 0.8 Mach and 35,000 ft (10,668 m), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned per hour divided by lbf of thrust the engine produces at that minimum point.
  • 'TSFC' Thrust Specific Fuel Consumption
  • Fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] ° '5 .
  • the "Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.52 m/second).
  • an example turbine blade 60 is illustrated, which may be suitable for the high pressure turbine 54, for example.
  • the turbine blade 60 is used in a first stage of the high pressure turbine 54, although the disclosed internal cooling features may be used for any blade or stator vane, or other components within a gas turbine engine.
  • the turbine blade 60 includes an airfoil 66 extending in a radial direction R from a platform 64, which is supported by a root 62, to a tip 68.
  • the airfoil 66 includes pressure and suction sides 74, 76 extending in the radial direction R and joined at a leading edge 70 and a trailing edge 72, as shown in Figures 2-4.
  • Figure 3 schematically illustrates the cooling passages 78 within the airfoil
  • the cooling passages 78 are supplied with cooling fluid F by first, second and third feed cavities 80, 82, 84.
  • the inlets to the feed cavities 80, 82, 84 are discrete from one another.
  • eight cooling passageways 86-100 are provided.
  • the first feed cavity 80 supplies cooling fluid F to first and second cooling passageways 86, 88.
  • the first and second cooling passageways 86, 88 are both connected directly to the first feed cavity 80 and terminate near the tip 68.
  • the second feed cavity 82 supplies cooling fluid F to a serpentine cooling passage provided by third, fourth and fifth cooling passageways 90, 92, 94.
  • the third feed cavity 84 supplies cooling fluid to sixth, seventh and eighth cooling passageways 96, 98,
  • the sixth cooling passageway 96 terminates in a tip flag near the tip 68.
  • the eighth cooling passageway 100 provides trailing edge cooling exits at the trailing edge 72 of the airfoil 66.
  • the cooling passageways 86-100 are spaced apart from one another in a chord- wise direction that extends from the leading edge 70 to the trailing edge 72.
  • the cooling passageways 86-100 extend in the radial direction R from the platform 64 to the tip 68.
  • the airfoil 66 can be divided from a design standpoint into first, second and third span regions 102, 104, 106.
  • the first span region 102 is provided near the platform 64
  • the third span region 106 is provided near the tip 68.
  • the second span region 104 is arranged radially between the first and third span regions 102, 106.
  • the span extends in the radial direction from a 0% span near the platform 64 to a 100% span near the tip 68.
  • the first span region extends 0-20% the span +1-5%
  • the second span region extends 20-60% the span +1-5%
  • the third span region 106 extends 60-100% the span +1-5%.
  • the airfoil 66 is provided by spaced apart pressure and suction walls 83, 85 that are joined at leading and trailing edges 70, 72. Intermediate walls 81 interconnect the pressure and suction walls 83, 85 to provide the cooling passages 78.
  • Each of the cooling passageways 86-100 include interior pressure and suction surfaces
  • Chevron trips strips may be used in conjunction with alternate trip strip configurations (i.e. skewed, normal, segmented, etc.) within a given cooling flow circuit or passage. Management of varying radial and axial external heat load along the airfoil surface- is necessary to achieve aggressive turbine performance, cooling flow, and life goals. Sequence of trip strip configuration array types (chevron, skewed, normal etc.) within a cooling circuit or cooling passage may vary enabling more flexibility and tailoring of unique trip strip array arrangements to address local co vective cooling requirements. Chevron trip strips may be used in the highest heat load area and then transition to skewed trip strips in lower heat load locations better balancing local metal temperatures in said region, resulting in lower thermal strains and improved oxidation and thermal mechanical fatigue capability.
  • a first trip strip geometry 108 is used to provide first heat transfer and pressure drop characteristics.
  • a second trip strip geometry 110 corresponding to a chevron-shape provides second heat transfer and pressure drop characteristics that are different than those provided by the first trip strip geometry 108.
  • the first trip strip geometry 108 is provided by a single linear trip strip skewed relative to the flow direction through the respective cooling passageway.
  • chevron-shaped trip strips 110 are provided in at least some of the cooling passageways 86-100.
  • the chevron trip strips 1 10 include protrusions 112 arranged in a V-shape forming an apex 1 14.
  • the apex 114 faces the flow direction.
  • the use of chevron trip strips 1 10 reduces the possibility of airfoil cracking by maximizing heat transfer and reducing temperatures at the expense of pressure drop in the locations 0%-30% span. This is achieved by employing chevron trip strips along at least some locations from 0% to 60% span in all three feed cavities 102-106.
  • the individual segment length of the chevron trip strip protrusions may vary, depending on die desired flow characteristics of the chevron, and may be unequal, as shown in Figure 7. Where required, non uniform segment lengths result in different boundary layer growth along each unique segment length of the chevron trip strip, resulting in more or less local convective heat transfer. The non-uniformity in local convective heat transfer enables the local thermal cooling effectiveness to be tailored and distributed to improve local internal convective heat transfer where required to offset or mitigate regions of high local heat load.
  • the midbody feed cavity 82 has the chevron trip strips 110 extending from 0% to 100% span in the fifth cooling passageway 94.
  • the use of chevron trip strips provide an additional 20%-25% of heat transfer for same amount of pressure drop when compared to using skewed trip strips only. This configuration allows a greater temperature reduction in the locations of the chevron trip strips, primarily in the 0%-3()% span area, which may be susceptible to cracking.
  • Chevron trip strips 1 10 are provided from 0% to 60% span +1-5% in the second and sixth cooling passageways 88, 96.
  • the trip strips 108, 1 10 are designed to have a height of 0.015 inch (0.38mm) +/-0.002 inch (0.05 mm) and a P/E of 6.7 +/-0.3, while the trailing edge cavity 84 has a height of 0.012 inch (0.30mm) +/-0.002 inch (0.05 mm) and a P/E of 8.3 +/-0.3.
  • trip strip heights of 0.007 - 0.020 inch (0.18 - 0.51 mm) and P/E ratios of 4-12 may be used.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une surface portante d'un moteur à turbine à gaz qui comporte des parois espacées de pression et d'aspiration réunies au niveau de bords d'attaque et de fuite pour créer une surface portante. Des parois intermédiaires relient entre elles les parois de pression et d'aspiration pour fournir des passages de refroidissement. Les passages de refroidissement ont des surfaces de pression et d'aspiration intérieures qui sont chacune disposées sur les parois de pression et d'aspiration. Des barrettes perturbatrices comportent une barrette perturbatrice en forme de chevron qui est disposée sur au moins l'une des surfaces de pression et d'aspiration intérieures.
PCT/US2013/051796 2012-08-22 2013-07-24 Éléments de refroidissement interne de surface portante de moteur à turbine à gaz WO2014031275A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP13830772.3A EP2888462B1 (fr) 2012-08-22 2013-07-24 Éléments de refroidissement interne de surface portante de moteur à turbine à gaz

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/591,773 2012-08-22
US13/591,773 US9157329B2 (en) 2012-08-22 2012-08-22 Gas turbine engine airfoil internal cooling features

Publications (1)

Publication Number Publication Date
WO2014031275A1 true WO2014031275A1 (fr) 2014-02-27

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US (1) US9157329B2 (fr)
EP (1) EP2888462B1 (fr)
WO (1) WO2014031275A1 (fr)

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US10174620B2 (en) 2015-10-15 2019-01-08 General Electric Company Turbine blade
US10208605B2 (en) 2015-10-15 2019-02-19 General Electric Company Turbine blade
US10370978B2 (en) 2015-10-15 2019-08-06 General Electric Company Turbine blade
US10443398B2 (en) 2015-10-15 2019-10-15 General Electric Company Turbine blade

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US10156157B2 (en) * 2015-02-13 2018-12-18 United Technologies Corporation S-shaped trip strips in internally cooled components
US10422233B2 (en) 2015-12-07 2019-09-24 United Technologies Corporation Baffle insert for a gas turbine engine component and component with baffle insert
US10280841B2 (en) 2015-12-07 2019-05-07 United Technologies Corporation Baffle insert for a gas turbine engine component and method of cooling
US10577947B2 (en) * 2015-12-07 2020-03-03 United Technologies Corporation Baffle insert for a gas turbine engine component
US10337334B2 (en) 2015-12-07 2019-07-02 United Technologies Corporation Gas turbine engine component with a baffle insert
US10196904B2 (en) 2016-01-24 2019-02-05 Rolls-Royce North American Technologies Inc. Turbine endwall and tip cooling for dual wall airfoils
US10590778B2 (en) 2017-08-03 2020-03-17 General Electric Company Engine component with non-uniform chevron pins
US10837291B2 (en) 2017-11-17 2020-11-17 General Electric Company Turbine engine with component having a cooled tip
US11002138B2 (en) * 2017-12-13 2021-05-11 Solar Turbines Incorporated Turbine blade cooling system with lower turning vane bank
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EP2888462A4 (fr) 2016-08-10
EP2888462A1 (fr) 2015-07-01

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